This invention relates to gas turbines and, more particularly, to a concept for efficiently reducing the temperature of air used to cool high temperature turbine rotor blades.
It is well understood that gas turbine engine shaft horsepower and specific fuel consumption (which is the rate of fuel consumption per unit of power output) can be improved by increasing turbine inlet temperatures. However, current turbines are limited in inlet temperature by the physical properties of their materials. To permit turbines to operate at gas stream temperatures which are higher than the materials can normally tolerate, considerable effort has been devoted to the development of sophisticated methods of turbine cooling. In early gas turbine engine designs, cooling of high temperature components was limited to transferring heat to lower temperature parts by the method of conduction, and air-cooling technology was limited to passing relatively cool air across the face of the turbine rotor disks. In order to take advantage of the potential performance improvements associated with higher turbine inlet temperatures, modern turbine cooling technology utilizes air-cooled hollow turbine nozzle vanes and blades to permit operation at inlet gas temperatures in excess of 2000.degree. F. (1094.degree. C.). Various techniques have been devised to cool these hollow blades and vanes. These incorporate two basic forms of air cooling, either singly or in combination, depending upon the level of gas temperatures encountered and the degree of sophistication permissible. These basic forms of air cooling are known as convection and film cooling. U.S. Pat. Nos. 3,700,348 and 3,715,170, assigned to the assignee of the present invention, are excellent examples of advanced turbine air-cooling technology incorporating these basic air-cooling forms. However, the benefits obtained from sophisticated air-cooling techniques are at least partially offset by the extraction of the necessary cooling air from the propulsive cycle. For example, probably the most popular turbine coolant today is air which is bled off the compressor portion of the gas turbine engine and is routed to the hollow interior of the turbine blades. The compressor air, having a temperature much less than that of the turbine flow path gas stream, absorbs heat from the turbine blades to maintain the blades at an acceptable temperature. When this heated cooling air leaves the turbine blades, perhaps as a coolant film, this heat energy is lost to the propulsive cycle since the cooling air is normally mixed with the exhaust gases and ejected from an engine nozzle. More particularly, the air that is bled from the compressor and used as cooling air for the turbine rotor blades has had work done on it by the compressor. However, because it is normally reintroduced into the flow path gas stream downstream of the turbine nozzle, it does not return its full measure of work to the cycle as it expands through the turbine. Additionally, the reintroduction of cooling air into the gas stream produces a loss in gas stream total pressure. This is a result of the momentum mixing losses associated with injecting a relatively low total pressure cooling air into a high total pressure gas stream. The greater the amount of cooling air which is routed through the turbine blades, the greater the losses become on the propulsive cycle. Thus, while turbine blade cooling has inherent advantages, it also has associated therewith certain inherent disadvantages which are functions of the quantity of cooling air used in cooling the turbine rotor blades.
It will, therefore, be appreciated that engine performance can be increased by reducing the amount of cooling air required by the turbine rotor blades.